Higher operating temperatures for gas turbine engines are continuously sought in order to increase their efficiency. However, as operating temperatures increase, the high temperature durability of the components of the engine must correspondingly increase. Significant advances in high temperature capabilities have been achieved through the formulation of nickel and cobalt-base superalloys, though such alloys alone are often inadequate to form components located in certain sections of a gas turbine engine, such as the turbine, combustor and augmentor. A common solution is to thermally insulate such components in order to minimize their service temperatures. For this purpose, thermal barrier coating (TBC) systems formed on the exposed surfaces of high temperature components have found wide use.
To be effective, thermal barrier coatings must have low thermal conductivity, strongly adhere to the article, and remain adherent throughout many heating and cooling cycles. The latter requirement is particular demanding due to the different coefficients of thermal expansion between materials having low thermal conductivity and superalloy materials typically used to form turbine engine components. Thermal barrier coatings capable of satisfying the above requirements have generally included a metallic bond coat deposited on the component surface, followed by an adherent ceramic layer that serves to thermally insulate the component. In order to promote the adhesion of the ceramic layer to the component and prevent oxidation of the underlying superalloy, the bond coat is typically formed from an oxidation-resistant alloy such as MCrAlY where M is iron, cobalt and/or nickel, or from an oxidation-resistant intermetallic such as a diffusion aluminide or platinum aluminide. Various ceramic materials have been employed as the ceramic layer, particularly zirconia (ZrO.sub.2) stabilized by yttria (Y.sub.2 O.sub.3), magnesia (MgO) or another oxide. These particular materials are widely employed in the art because they can be readily deposited by plasma spray, flame spray and vapor deposition techniques, and are reflective to infrared radiation so as to minimize the absorption of radiated heat.
A significant challenge of thermal barrier coating systems has been the formation of a more adherent ceramic layer that is less susceptible to spalling when subjected to thermal cycling. For this purpose, the prior art has proposed various coating systems, with considerable emphasis on ceramic layers having enhanced strain tolerance as a result of the presence of porosity, microcracks and segmentation of the ceramic layer. Microcracks generally denote random internal discontinuities within the ceramic layer, while segmentation indicates the presence of microcracks or crystalline boundaries that extend perpendicularly through the thickness of the ceramic layer, thereby imparting a columnar grain structure to the ceramic layer. Ceramic layers for thermal barrier coating systems employed in high temperature applications of a gas turbine engine are typically deposited by electron beam physical vapor deposition (EBPVD) techniques that yield the desirable columnar grain structure, which is able to expand without causing damaging stresses that lead to spallation. A strong adherent continuous oxide surface layer is often formed over the bond coat to protect the bond coat from oxidation and hot corrosion, and to provide a firm foundation for the PVD columnar ceramic layer.
Though significant advances have been made in attaining spallation-resistant thermal barrier coatings, there is the inevitable requirement to repair such coatings under certain circumstances. Typically, spallation occurs in localized regions or patches, either during engine service or during post coating processing of the coated component. In addition to service-related incidents, processing flaws that occur during the formation of the thermal barrier coating system can also necessitate repairs. One example is during the deposition of an EBPVD ceramic layer, where a molten liquid splat of ceramic may mar the coating, rendering the coating unacceptable for service.
The current state-of-the-art repair methods often result in removal of the entire thermal barrier coating system, i.e., both the ceramic layer and bond coat, after which the bond coat and ceramic layer must be redeposited. Due to a heightened resistance to spallation, a columnar ceramic layer is very difficult to remove. Prior art techniques for removing thermal barrier coatings have generally involved grit blasting or subjecting the coating to an alkaline solution at high temperatures and pressures. Grit blasting is a slow, labor-intensive process and erodes the ceramic layer and bond coat, as well as the substrate surface beneath the coating. With repetitive use, the grit blasting process eventually destroys the component. The use of an alkaline solution to remove a thermal barrier coating requires the use of an autoclave operating at high temperatures and pressures, and also results in removal of the entire coating system.
A more recent process for repairing a thermal barrier coating system is disclosed in U.S. patent application Ser. No. 08/362,377 to Reeves et al., assigned to the assignee of this invention. The disclosed process entails heating the thermal barrier coating to about 870.degree. C. or more while exposing the coating to a halogen-containing powder, which causes the coating to deteriorate to the extent that it separates from the underlying substrate. In doing so, both the ceramic layer and the bond coat must be redeposited as a result of the halogen-containing powder attacking each of these layers.
From the above, it can be appreciated that a process for repairing a thermal barrier coating system that does not damage or remove the bond coat would be advantageous. In particular, if a process were available that attacked only the ceramic layer, the labor, processing and costs required to refurbish a thermal barrier coating system could be significantly reduced.